Supersonic combustion nozzle



June24, 1969 R. o. BOE

- SUPERSONIC COMBUSTION NOZZLE Sheet Filed July '26, 1966 /PO, A/A/ 0.505,

INVEN'IOR.

United States Patent M 3,451,221 I SUPERSONIC COMBUSTION NOZZLE Rollin0. Boe, Canoga Park, Califi, assignor to The Marquardt Corporation, VanNuys, Calif., a corporation of California Filed July 26, 1966, Ser. No.568,029 Int. Cl. F02g 3/00; F021: 9/00, 1/00 US. Cl. 60-258 14 ClaimsABSTRACT OF THE DISCLOSURE A convergent-divergent supersonic combustionnozzle in which high pressure oxidizer is introduced subsonically to theconvergent portion for expansion to supersonic velocity and fuel isintroduced in the vicinity of the throat parallel to the oxidizer flowfor supersonic combustion therewith in the divergent portion. The fueland oxidizer may be hypergolic, or combustion may be initiated byelectric or pilot burner means.

This invention relates to a supersonic combustion nozzle and moreparticularly to a nozzle having burning in the supersonic portion of thenozzle thereby eliminating heating problems in the region of the sonicthroat.

In present conventional rockets, combustion takes place in a subsonicchamber and the resulting hot gas flow is exhausted through a sonicthroat and a supersonic nozzle. The heat transfer to the nozzle ismaximum in the region of the sonic throat since the heat transfer islargest where the diameter is minimum. The high temperatures encounteredwith high Mach number nozzles have resulted in a serious heating problemof the material used to make the sonic throat. In order to keep thethroat and the subsonic combustion chamber cool, water jackets and coolair boundary layers must be utilized. These methods of cooling take heatfrom the main gas flow, resulting in reduced efficiency.

The supersonic combustion nozzle of the present invention utilizesburning in the supersonic stream of the nozzle in order to reduce thecooling requirement in the critical sonic portion of the nozzle and toreduce the size of the sonic throat. Also, the need for large subsoniccombustion chambers is eliminated and most of the heat of combustionremains in the supersonic exhaust flow. The nozzles have applicationboth as propulsion nozzles and as flight simulation test nozzles. Anelectrical arc can be utilized to accomplish ignition of the fuel andoxidizer in the supersonic region or a small pilot burner located in thesupersonic region can be utilized. Also, the fuel can be introduceddirectly into the supersonic oxidizer flow stream for auto ignition whencompatible fuel and oxidizers are utilized. In most cases, the fuelinjection system will introduce fuel in the vicinity of the sonic throatand the fuel is added to the oxidizer in the sonic region either throughnozzles in the flow stream line or through the walls of the nozzle. Thenozzle configuration for supersonic combustion differs from the standardnozzle configuration since large expansion will take place duringsupersonic combustion requiring larger expansion ratios than encounteredin normal supersonic nozzles.

It is therefore an object of the invention to provide a supersoniccombustion nozzle in which combustion occurs in the supersonic portionof the nozzle thereby eliminating heating problems in the region of thesonic throat.

Another object of the present invention is to provide a supersoniccombustion nozzle in which the fuel and oxidizer are mixed together andcombusted in the supersonic flow portion of the nozzle, thereby reducingthe cooling requirements in the sonic portion of the nozzle 3,451,221Patented June 24, 1969 and eliminating the need for large subsoniccombustion chambers.

Another object of the invention is to provide a supersonic combustionnozzle in which fuel and oxidizer are mixed in the vicinity of thesupersonic portion of the nozzle and are combusted by ignition meanslocated in this region.

These and other objects of the invention not specifically set forthabove will become readily apparent from the accompanying description anddrawings in which:

FIGURE 1 is a section through a supersonic combustion nozzle of thepresent invention showing the air and fuel manifolds;

FIGURE 2 is an enlarged transverse section along line 22 of FIGURE 1showing the fuel tubes within the nozzle throat;

FIGURE 3 is a section of a modified supersonic combustion nozzle inwhich the oxidizer is introduced to indi- Vidual tubes, each containinga fuel supply tube;

FIGURE 4 is a transverse section along line 4-4 of FIGURE 3 showing theoxidizer manifold;

FIGURE 5 is a section of a third modification of the supersoniccombustion nozzle showing the fuel tubes inserted through the sides ofthe nozzle and the electrodes located in the supersonic flow region;

FIGURE 6 is a transverse section along line 66 of FIGURE 5;

FIGURE 7 is a sectional view of a fourth modification of the supersoniccombustion nozzle in which the fuel and oxidizer flow through separatenozzles and auto-ignite in the supersonic region;

FIGURE 8 is a transverse section along line 8-8 of FIGURE 7 showing themanifolding for the fuel; and

FIGURE 9 is a schematic of the nozzle showing the various regions of thenozzle.

Referring to FIGURES 1 and 2, the supersonic combustion nozzle comprisesa nozzle block 10 having an entrance '11 leading to the nozzle throat12. The diverging portion 13 of the nozzle is connected at end 13a tothe nozzle block 10 and the other end 13b carries a flange 14. A cover15 extends from the nozzle block 10 parallel to the diverging portion 13and has a step section 16. Spiral spacer rings 17 are connected to thediverging nozzle portion 13 to form an annular cooling passage 17a between the diverging nozzle portion 13 and the cover 15. A cylinder -18projects from the flange 14 and has an end 18:: received in an annularcup 19 containing a seal 20 in the form of a flexible gasket. The cup 19is secured to a larger annular member 21 and a plurality of bolts 22project between the member 21 and a circular flange 23 carried by thecylinder 18. Tightening of the bolts 22 will cause the end 1811 tosqueeze the gasket 20 into sealing relationship with the step section 16while still permitting relative movement of the cylinder 18 and thesection 16. Since the spacer rings 17 are secured only to the divergingnozzle portion 13 and since the gasket 20 can move relative to theoffset cover section 16, it is apparent that a temperature differentialwill not produce stresses in this nozzle structure.

A casing member 30 has a flange 31 which is connected to a flange 32 onthe nozzle block 10 by means of a plurality of bolts 33. The flange 31is located at the end of a cylindrical portion 35 of the casing memberand portion 35 connects with a conical portion 36 leading to a secondcylindrical portion 37. A pilot combustion chamber 40 is located withinthe casing member 30 and is spaced therefrom by means of spacer rings 41connected only to the chamber 40. A base plate 43 closes the end ofcasing member 30 and also supports and closes the end of chamber 40. Thebase plate 43 supports an igniter 44 of any well known standardconstruction and fuel passage 45 and oxidizer passage 46 dischargeadjacent the igniter in order to produce combustion within the pilotcombustion chamber. The reduced discharge end 47 of combustion chamber40 passes through the nozzle 12 and opens into the region of the nozzlewhere the flow is supersonic in order to provide a small pilot burner. Asmall amount of fuel and oxidizer burn subsonically in a combustionchamber 40 and the combustion products flow through the pilot burner 47at a high enough temperature to ignite the cold oxidizer and fuel flowsto the main nozzle.

A circular manifold 50 surrounds the cylindrical portion 35 of easingmember 30 and connects to an entrance passage 51 which supplies highpressure oxidizer, such as air or oxygen, to the manifold. The manifoldhas four radial passages 52 (only two of which are shown) which lead toa space 53 located between the nozzle block and a partition 54 in casingportion 35. The high pressure oxidizer supplied to space 53 flowsthrough the nozzle throat 12 and obtains a supersonic velocity in theregion of 55 downstream of the throat. A smaller circular fuel manifold56 also surrounds the cylindrical casing portion 35 and twelve radialpassages 57 lead from the manifold through the casing portion 35 andproject through the nozzle throat 12 so that the end 57a of each passageis located adjacent the end of pilot burner 47. Fuel is supplied tomanifold 56 through the passage 60 leading to a fuel supply source.

In operation of the nozzle, oxidizer at high pressure flows frommanifold 51 through the space 53 and through nozzle throat 12 and hassupersonic velocity in the nozzle region 55. However, the temperatureand static pressure of the oxidizer have been considerably reduced atregion 55. For instance, air introduced at about 400 psi. into themanifold 50 would be at about 50 p.s.i. in the region 55 and at atemperature of -170 F. However, the air has reached a velocity of about1700 feet per second corresponding to Mach 2. The static pressure of thefuel which is introduced into the supersonic region 55 through the fueltubes 57 will be substantially the same as the static pressure of theoxidizer in region 55. In order to prevent shock waves from building upwhen the hot flow from the pilot combustion chamber 40 meets thesupersonic cold flow, the velocity and static pressure of the hot flowfrom chamber 40 is made substantially the same as the velocity andstatic pressure of the supersonic cold flow. This is accomplished byhaving the hot flow enter the region 55 at a subsonic velocity (aboutMach 0.5) and at a temperature (about 5000 F.) which will provide thesame static pressure and velocity as the cold flow.

Since the pilot combustion chamber operates at a low combustionpressure, the thick walled chambers are not required to contain thepressure. Also, the ratio between the oxidizer flow and the hot flowfrom the combustion chamber 40 can be in the neighborhood of 50 to 1.Therefore, high volumes of flow from the combustion chamber 40 are notrequired. Since the pilot burner tube 47 of the pilot combustion chamberis parallel to the supersonic cold flow through the nozzle, reduction ofshock waves in the supersonic flow is obtained and more uniformcombustion results from starting the burning at the center of the flow,A schematic illustration of the various regions in the nozzle isillustrated in FIGURE 9. Subsonic flow exists in region 1 upstream ofthe throat 12 and the first expansion of the oxidizer to a predeterminedMach number for burning will occur in the region 2. In region 3,constant Mach number combustion will occur and in region 4, the finalexpansion to the maximum Mach number will result.

A modification of the supersonic combustion nozzle is illustrated inFIGURES 3 and 4. The nozzle has throat portion 65 connected with adiverging section 66. The entrance end 67 to the nozzle is connected bymeans of a scalloped band 64 to the ends 68a of a plurality of circularoxidizer tubes 68 which are welded together at their abutment location69 to form a circular configuration. The entrance ends 68b of the tubes68 are connected to an annular space 70 defined by sides 72 and 73 andan annular base 74. An annular oxidizer manifold 75 surrounds the tubes68 and is connected by passage 76 to a high pressure oxidizer supply.Four radial passages 78 connect manifold 75 to space 70. Thus, oxidizerfrom manifold 75 is conducted through the twelve tubes 68 to theentrance of the nozzle. The oxidizer then passes through the nozzlethroat 65 and is supersonic in the region 80.

The interior surface of the tubes 68 form the interior surface of pilotcombustion chamber 81 which produces a hot gas flow at the end 83 forigniting the main supply of fuel and oxidizer. The pilot burner end 83of the combustion chamber 81 is attached to the interior of the ends 68aof tubes 68 by a scalloped band 83a (not shown) and projects through thethroat 65 of the nozzle into the supersonic cold flow. An igniter 84 ofstandard construction is located in the back wall 85 which closes theend of chamber 81 and fuel line 86 and oxidant line 87 terminate in thevicinity of the igniter in order to produce hot gas by combustion. Thesmall flow of hot gas at the exit end 83 is of sufficiently hightemperature to ignite the main oxidant and fuel supply to the nozzle. Afuel tube 91 extends through the center of each of the oxidizer tubes 68and the ends 91a of the tubes 91 are located adjacent the end 83 of thepilot burner. The other ends 91b of the fuel tubes extend through side71 and connect with an annular fuel manifold 95 which is connected to asource of fuel supply through the passage 96.

The operation of the modification of FIGURES 3 and 4 is similar to thatof the first embodiment since high pressure oxidizer is introduced toeach of the tubes 68 and passes through the nozzle throat 65 to becomesupersonic in the region 80. At this location, hot gas flow isintroduced from the pilot combustion chamber 81, and also fuel isintroduced through the individual tubes 91. The resulting mixture offuel and oxidizer is ignited by the hot gas fiow in the supersonicregion 80. In this modification, the individual air passages 68 areconnected together to provide an annular combustion chamber 81 withoutadditional structure so that a light nozzle motor is provided. Theoxidizer flow through tubes 68 cools the portion of the tubes 68 whichform the interior surface of the pilot combustion chamber 81. Also, theexterior surface of the tubes 68 provide the outer surface of the nozzleand additional cowlings and structure are not required Referring to athird modification shown in FIGURES 5 and 6, the nozzle has an entranceportion and a diverging portion 101 connected together by the throatportion 102. A plurality of fuel tubes 103 extend through the entranceend wall of the nozzle and their ends 103a terminate in the vicinity ofthe throat of the nozzle. The other end 1031; of each tube connects witha suitable fuel manifold (not shown) and an oxidizer is introduced underhigh pressure into the entrance portion 100 of the nozzle from anoxidizer source. A pair of electrodes 104 and 105 extend through thethroat of the nozzle to produce an electric are 106 within thesupersonic region 107 of the nozzle. The electric arc is utilized toignite mixtures of fuel and oxidizer which are not self-igniting. Sincethe electrodes are inserted parallel to the flow, they produce minimumeffect on the supersonic air stream and the supersonic flow will serveto cool the electrodes. Two alternate locations for the ends 103a aredesignated by the positions A and B in FIGURE 5. In the case of fuelsand oxidizers which can be mixed without auto-ignition, the ends of thetubes can be in the subsonic region ahead of the sonic throat asindicated by position A of the nozzle ends. However, for manyfuel-oxidizer combinations, this is not possible and therefore thenozzle ends must be located at least as far downstream as position B tointroduce the fuel in the vicinity of the sonic throat. When the fuelnozzles are located in position A upstream of sonic throat, they allowthe maximum time for the fuel and oxidizer to mix.

In the case of fuel-oxidizer combinations that autoignite, the etfectivearea of the sonic throat can be reduced to the extent that combustiontakes place ahead of the sonic throat by placing the nozzle exit endsupstream of the throat. Thus, the effective area of the nozzle can bechanged by moving the fuel nozzle to vary the amount of combustion whichtakes place before reaching the throat. The change in efiective nozzlearea can result in a change of the effective area ratio between thesonic throat and the supersonic exit nozzle, thereby serving as a meansfor varying Mach number of a fixed configura tion nozzle. The Machnumber can also be varied by changing the fuel-oxidizer ratio since theamount of expansion due to supersonic combustion will vary. Generally,the Mach number goes down as the fuel oxidizer ratio increases. Thus,the third modification of the invention provides a convenient structurein which to vary the location of the outlet ends of the fuel tubes inorder to provide a large variety of conditions for the fueloxidizercombustion and this embodiment can also employ electric arc type ofigniter.

The fourth embodiment of the invention is illustrated in FIGURES 7 and 8wherein the fuel and oxidizer are introduced through individual annularnozzles into supersonic layers of mixture which will auto-ignite. Thenozzle consists of a divergent portion 115 which is connected by throatportion 114 with an entrance portion 116 through which the oxidizerflows. Four annular sleeves 117 are located in the entrance portion andeach sleeve terminates in an enlarged end 117a to form a nozzle 119 withits adjacent sleeve end 117a and a nozzle 119a with throat 114. Theindividual sleeves are attached to four passages 130 leading from a fuelmanifold passage 131 which extends transversely across the entranceportion 116. The outer annular passage 120 receives oxidizer fromentrance 116 and discharges the oxidizer through nozzle 119a. Thepassages 121 and 122 also receive oxidizer and each passage dischargesthrough a nozzle 119. The annular passages 125 and 126 each receive fuelfrom two passages 130 and passages 125 and 126 each dis charge through anozzle 119. The passages 125 and 126 are blocked from the oxidizer inentrance portion 116 by the back plates 133 and 134 which containopenings for passages 130. In the fourth modification, both the fuel andoxidizer are discharged in annular layers and at supersonic velocitiesinto the divergent section of the nozzle. Fuels and oxidizers areutilized that will autoignite upon contact without additional ignitionmeans and the alternate positioning of the layers of fuel and oxygenwill facilitate the auto-ignition. It is understood that the number offuel and oxidizer nozzles can be varied in order to obtain the properflow of these substances to provide the desired fuel-oxidizer ratio.

By the present invention there is provided a supersonic combustionnozzle in which high pressure, cold flow passes through a sonic throatto reach supersonic velocity and the cold flow is thereafter mixed withthe fuel so that combustion can take place in the supersonic region ofthe nozzle. Ignition can be initiated in the supersonic region eitherelectrically or by means of a hot gas from a pilot burner. While it isusually desirable to introduce the fuel into the supersonic region ofcold flow, the point of introduction of the fuel can be varied asindicated to produce various modifications of nozzle performance, suchas permitting a small amount of the combustion to take place in thesubsonic region and thereby vary the effective area of the sonic throat.Since substantially all combustion takes place in the supersonic region,only a small amount of cooling is required in the throat portion of thenozzle and no large subsonic combustion chamber is required. In general,supersonic combustion will require larger expansion ratios thanencountered in present subsonic nozzles. Also, the length of thesupersonic combustion chamber can be minimized by burning at lowsupersonic Mach numbers.

Various other modifications in addition to those described herein arecontemplated by those skilled in the art without departing from thespirit and scope of the invention hereinafter defined by the appendedclaims.

What is claimed is:

1. A supersonic combustion nozzle comprising;

a nozzle throat portion connecting with a convergent nozzle entranceportion and discharging through a divergent nozzle portion;

means for introducing high pressure oxidizer subsonically to saidentrance portion for expansion through said throat portion to reachsupersonic velocity in said divergent portion; and

means for introducing fuel into said nozzle substantially parallel tothe oxidizer flow in the vicinity of said throat portion to producesupersonic combustion in the divergent portion of the nozzle.

2. A supersonic combustion nozzle as defined in claim 1 wherein saidfuel introducing means comprises a plurality of fuel tubes having theiroutlet in the supersonic flow region of the nozzle.

3. A supersonic combustion nozzle as defined in claim 2 having ignitionmeans located in the vicinity of the outlets of said fuel tubes in saidsupersonic region for initiating combustion downstream of the nozzlethroat portion.

4. A supersonic combustion nozzle as defined in claim 3 wherein saidignition means comprises a subsonic combustion pilot burner to produce ahot flow for igniting said fuel and oxidizer.

5. A supersonic combustion nozzle as defined in claim 4 wherein saidpilot burner comprises a tube axially positioned at the center of saidthroat portion and extending upstream of said throat to connect withsaid pilot combustion chamber.

6. A supersonic combustion nozzle as defined in claim 3 wherein saidignition means comprises electrodes eX- tending axially of said throatportion and terminating in the vicinity of the supersonic oxidizer flow.

7. A supersonic combustion nozzle as defined in claim 1 wherein saidfuel introducing means comprises a plurality of fuel tubes having theiroutlets slightly upstream of said throat portion to allow maximum timefor fuel and oxidizer to mix, said fuel and oxidizer being of the typewhich do not auto-ignite.

8. A supersonic combustion nozzle as defined in claim 1 wherein saidfuel introducing means introduces said fuel slightly upstream of saidthroat to vary the elfective area of the nozzle; said fuel and oxidizerbeing of the type which auto-ignite.

9. A supersonic nozzle as defined in claim 4 wherein said oxidizerintroducing means comprises a plurality of circular tubes attachedtogether in a circular configuration and discharging into said throatportion of said nozzle, the interior surfaces of said tubes forming saidpilot combustion chamber, said fuel introducing means comprisingindividual fuel tubes extending through each of said oxidizer tubes.

10. A supersonic combustion nozzle .as defined in claim 4 wherein saidthroat portion comprises a plurality of individual nozzles discharginginto said divergent portion, said oxidizer and fuel introducing meanscomprising a plurality of passages connecting some of said nozzles tooxidizer and other of said nozzles to fuel.

11. A supersonic combustion nozzle as defined in claim 2 wherein saidfuel tubes extend through said throat portion.

12. A supersonic combustion nozzle as defined in claim 4 wherein saidoxidizer introducing means introduces said oxidizer at a pressure toproduce supersonic flow downstream of said throat portion at a reducedstatic pressure, said pilot burner producing a hot flame at a velocityand static pressure matching that of supersonic oxidizer in order toprevent the development of shock waves in the region of mixing of thefuel and oxidant.

13. A supersonic combustion nozzle as defined in claim 2 wherein saidfuel introducing means introduces said fuel to said oxidizer at thestatic pressure of said oxidizer.

14. A supersonic combustion nozzle as defined in claim 1 having ignitionmeans located in the region of supersonic velocity.

References Cited UNITED STATES PATENTS 2,992,527 7/1961 Masnik 60-2703,095,694 7/1963 Walter 60261 3,112,988 12/1963 Coldren 60-270 9/1964Rocca 60261 10/1966 Sippel 60261 10/1966 Dugger 60-270 5/1967 Walter60-261 6/ 1967 Sanger 60270 8/1967 Rhodes 60270 FOREIGN PATENTS 1/1966Germany.

US. Cl. X.R.

